Independent landing monitoring pulse radar system

ABSTRACT

An Independent Landing Monitoring pulse radar system supplying angular information similar to these supplied by an I.L.S. system comprising circuits for improving the precision of the supplied data through compensation of the dynamic computation errors due to the aircraft displacement between successive data measurements, and circuits supplying additional position information such as location, velocity distance and estimated stopping distance. The system may also be operated for synthetic runway display, as a weather and as an anti-collision radar system.

United States Patent Gendreu et al.

[ INDEPENDENT LANDING MONITORING PULSE RADAR SYSTEM Inventors: RobertGendreu; Jean Genuist, both of Paris, France [73] Assignee: Thomson-CSF,Paris, France [22] Filed: Jan. 11, 1972 [30] Foreign ApplicationPriority Data Feb. 4, 1971 France ..7lO3732 21 Appl. No.: 217,071

US. Cl 343/5 LS, 343/5 W Int. Cl. G015 9/00 Field of Search 343/5 LS, 7TA, 5 W

References Cited UNITED STATES PATENTS 8/1947 Hopgood 343/5 LS X 8/1964Shellet 1 343/112 R 4/1965 Case et al 343/5 LS X Nov. 27, I973 3.l8l,l534/1965 Cella 343/5 LS X Primary Examiner-T. H. Tubbesing Attorney-JohnW. Malley et al.

[57] ABSTRACT An Independent Landing Monitoring pulse radar systemsupplying angular information similar to these supplied by an l.L.S.system comprising circuits for improving the precision of the supplieddata through compensation of the dynamic computation errors due to theaircraft displacement between successive data measurements, and circuitssupplying additional position information such as location, velocitydistance and estimated stopping distance. The system may also beoperated for synthetic runway display, as a weather and as ananti-collision radar system.

8 Claims, 2 2 Drawing Figures A. G CO SHEET 05 0F 11 P m A": u H H H 8?H m x 3m 52! f $25; 5 Q N: Em 2 m2: M kxtmaL n w u v 1x 9 5 mm N M K :3w k 5:73 5X2 A i i m J M r ON [l N A J :W R I v 9 Jazz 5:; 252 a l Lmmwvw mi .2 w 5 mgTlpfimo 5:: E @QWQ W PATH SHEET 07 0F 11 LICOMPUTEDIFFERE INTEBRMO PiME NERRH COMPUTER COMPUYE ADUE PO SWION DETEEYOR lINDEPENDENT LANDING MONITORING PULSE RADAR SYSTEM Independent LandingMonitoring systems or ILM are known which are combined with wheather andground display radars.

The present invention relates to those, among such systems, whichcomprise pulse-modulated transmitter and associated receiver circuitscombined with computing circuits, elaborating azimuthal and elevationangles, from the radar information and from known data related to theposition of associated ground reflectors.

It is an object of this invention to improve ILM systems with the viewof providing accurate angular information similar to those supplied byconventional ILS systems and in addition of providing the pilot withlocation, velocity, distance and estimated stopping distanceinformation.

According to the invention there is provided an Independent LandingMonitoring pulse radar system for use in cooperation with threereflectors, and with an auxiliary on board system supplying flightparameters said system supplying bearing and elevation informationrelative to the position of the aircraft with respect to the runway onwhich it has to land, similar to those supplied by conventional I.L.S.systems, said radar system further comprising circuit means forcompensating for the dynamic errors in the computation of said bearinginformation.

For a better understanding of the invention and to show how the same maybe carried into effect, reference will be made to the attached figuresin which FIGS. 1 to 5 show the reference trihedrons employed in thefollowing description FIG. 6 illustrates the angular parameters usuallyprovided to the pilot of an aircraft for carrying out an automaticlanding.

FIG. 7 illustrates an example of a system of reflectors cooperating withan ILM system according to the invention FIGS. 8 to 11 illustratevarious input and output parameters of the system according to theinvention FIGS. 12a and 12b are block diagrams of an improved ILM systemaccording to the invention FIGS. 13, l5, l6, l9 and 21 are circuitdiagrams of particular blocks of FIG. 12

FIGS. l4, l8, and show input parameters of the system shown in FIGS. 12and 13; and

FIG. 17 shows an example of a display system associated with the ILMsystem according to the invention.

Similar references designate similar elements throughout the variousfigures.

First, for the sake of clarity the reference trirectangular trihedralsnecessary to an understanding of the invention will be defined. Thesetrihedrals are shown in FIGS. 1, 2 and 3 namely the aircraft trihedralAXYZ, the measuring trihedral Axyz and the stabilized trihedral Ax, y,z,respectively.

The aircraft trihedral AXYZ shown in FIG. 1 as its name suggests, isfixed in the aircraft which is assumed to be reduced to the phase centerof the antenna of the ILM. The axes AX, AY, AZ are the bocy axes theaxis AX is the longitudinal or roll axis of the aircraft the axis AZ,perpendicular to AX, is in the plane of symmetry of the aircraft, it isthe normal axis or yaw axis the axis AY is perpendicular to thepreceding axes and is the lateral or pitch axis The measuring trihedralAxyz, shown in FIGS. 2, is that to which reference will mostly be madehereafter it is a trihedral which is stabilized in pitch and roll, andis derived from the aircraft trihedral by two rotations. The first oneis by I relative to the longitudinal axis, where I is the lateralattitude angle, which is considered positive when the right wing of theaircraft is lower than the left wing. This rotation by I brings the axisAY to Ay i.e. on the horizontal (shown in FIG. 5) The second rotation isby 0 being the longitudinal attitude angle considered positive when theaircraft is tail down, which brings the axes AX and AZ to Ax and Azwhich are respectively horizontal and vertical (0 is shown in FIG. 4)The stabilized trihedral Ax y z shown in FIG. 3 is deduced from themeasuring trihedral Ayxz by a rotation through AC about the axis Az, theaxes Az and A2 being coincident AC represents the instantaneousvariation in the course, which is positive when the aircraft deviatesfrom its route by turning in the clockwise direction FIGS. 2 and 3 showthe relation between the courses C and Ca i.e. the angles the axes Axand Ax make with the north direction AN, the relation between thebearings G, and G and that between the angles of elevation S and S ofany point M relative to the two trihedrals,

Onehas:

C=Ca+AC G =G +AC FIGS. 4 and 5 set forth the angle of pitch 0 and theangle 1 or the longitudinal and lateral attitudes. The plane of FIG. 4is the vertical plane containing the axes Ax and AX and the plane ofFIG. 5 is the plane perpendicular to AX at A. FIGS. 6 and 7 show theaircraft A in final approach ready to land upon the runway RW.

The point B on the runway axis, is the theoretical point of touchdownwhich is located around 300 m from the runway threshold E the verticalplane W passing through the runway axis defines the vertical plane ofalignment, the localiser transmitter of the conventional instrumentlanding system or ILS system, if any,being located upon the runway axisat B beyond the far end runway threshold F.

The axis TB contained in the plane W and making the angle p with therunway axis, represents the theorical glide path, which is theintersection between the plane W and the theorical glide path plane. Theintersection CC of this plane with the runway passes through B and isperpendicular to the runway axis The conventional glide path transmitterof the ILS system is located upon this pivot line.

These various reference elements, are well known to those skilled in theart of aerial navigation,and all the details thereon can be found in avariety of text books such for example as the OACI standards, Avionicsand Navigation Systems, edited by KAYTON AND FRIED, and published byJOHN WILEY & SONS, 1969 Library of Congress, Card number 69-13 679, P528,a.nd followings. The plane W is the vertical plane, perpendicular tothe runway axis, containing the aircraft A at the instant considered (itis assumed, of course, that the aircraft is a point in space, the phasecentre in fact of the airborne radar antenna) and A is the point at thesame altitude Z as the aircraft,

which is contained simultaneously in the planes W and The parametersrequired for the guidance of the aircraft arg e deviation angles 'y and8 y A,B, BA, i.e. the angular deviation in the azimuthal plane and 8 =TBB A, i.e. the angular deviation in the elevational plane, seen from B A,being the normal projection of A on the plane W. Angle p is generally ofthe order of 2 to 3 and the deviation angles 7 and 8 should not exceed 2and respectively.

The position of the aircraft is fully defined by the coordinates X Y,,,Z, of A in the rectangular axes system B X Y Z where B X is carried bythe axis of the runway RW and B,,Y is perpendicular to B X the axis B Zbeing vertical and extending upwardly. These axes are shown in FIGS. 10and 11 the orientation of the measuring trihedral Axyz is fully definedby the angle All! made by the projection of the axis Ax on thehorizontal plane (FIG.10) with the axis B,,X,,. In order to simplify thedrawing, Ax designates in FIGS. 10 and 11 in fact the projection of Axrespectively on the horizontal plane containing the runway and on thevertical plane W passing through the axis of the runway. The orientationin elevation of the radiation axis of the antenna of the system isdefined by the angle 1; (FIGS.9 and 11) that the projection AFa of thisaxis on the plane W makes with Ax. The projection of the axis Ax of thestabilised trihedral is represented with the same references at Ax, FIG.10.

The figures have not been drawn to scale for purposes of clarity inparticular the angles 8, -y and All: and the co-ordinate Y, of the pointA are usually much smaller than shown.

The runway is equipped with a reflector assembly located so as to definethe runway axis and the vertical plane W. Preferably this assembly isformed by the three reflectors B, B and B shown in FIGS. 7, 10, 11, 14and 18, reflector B being located at B (FIG.6) and reflectors B, and Bbeing on a straight line with the touch down point B,, and beingsymmetrical with each other with respect to that point.

This reflector system cooperates in a known manner with the ILM system.To this end the reflectors can be active reflectors, retransmittingreceived signal at a different frequency, thus avoiding guidance errorsdue to interferences with echoes from the ground or from other targetssuch as other landing aircraft.

The aircraft A is equipped with a long-range radar system of therecurrent pulse type to this end, its wavelength is made relativelylong, for example in the order of 3 cm.

During cruising flight, the system operates as a weather or meteo radarit produces pulses of duration 7,, for example of the order of 5 1.5length,and scans through an azimuthal angle of around i 90 at a scanningrate in the order of 60/sec. The operator can select the elevationalangle of the antenna.

The carrier frequency of the received echoes is equal to thetransmission frequency fb to within the Doppler frequency which isdefined by the relative speed of the targets.

Conventional frequency changer, amplifier and detection circuits displayupon a screen the radar map of the disturbances. Known arrangements makeit possible to utilise the same circuits in order to display majorgeographical contours, such, for example, as coast lines.

During the landing the scanning of the antenna in the azimuthal plane,is accelerated (boosted for example to l20/sec.), and, at the same time,limited to a sweep range of i 25, the elevational angle of the antennabeing fixed.

The length of the transmitted pulses is shortened to 1- for exampleO.4p.s and the receiver circuits which will be described in more detailhereinafter with reference to FIG. 12, are now arranged to receive notthe echoes of frequency f", but the echoes of different frequenciesf,=fl, Af(orf Af) returned by the reflectors or beacons B, B and B whichare assumed to be active (frequency modulating reflectors returningsignals at frequency f The rolling and pitching movements of theaircraft are of course compensated for, either through displacing theantenna, or in the processing of the received signals.

The antenna of the system is designed for operating, during thereception, as a sum-difference monopulse antenna in elevation, theantenna either comprising two radiating apertures-one of which is notused when the system is operated as a weather radar or comprising asingle aperture associated with known multimode excitation means.

The angles g, g 3 (FIG. 8) of the axes AB, AB AB with the axis AX of theaircraft are known from the instantaneous position of the antenna at themoment the targets are detected.

These angles are determined to within some few milli-radians if the beamhas a three-degree aperture angle. They enable the angle y to be roughlycomputed The glide slope angle of the aircraft being small, thecomputation of 'y can be effected in the horizontal plane, shown inFIGS. 8 and 10, A being the projection of A onto said plane. As long asthe aircraft is far away from the runway, the conditions set out in theFigure are strictly correct and 'y r /L g g r being the distance A Bwhich can be equated to the distance A B measured by the radar, L beingthe distance between B, and B and g being equal to g, g j2 As theaircraft gets closer, this approximation ceases being valid because theazimuthal angle of the point B can no longer be equated to g, g /2 Thenthe relationship is being valid, where b is the distance 8 8, that is tosay the projection onto the runway axis of the distance B, B which isabout m (FIG.10).

The distance A,A (FIG. 6) being small in regard of A B (FIG. 6) theangular deviation 6 can be computed quite simply from the elevationdeviations AS, and A5 supplied by the monopulse output of the radar, asFIG. 9 shows p being the theoretical glide slope and B the elevationalangle of A, seen from 8,, the relationship between 8 and [3 is 8 B p inother words 8=r /L (AS A8 p r being the distance A,B which is similar tothe distance AB measured by the radar.

The computations of y and 8 according to the expressions (l and (2)which are simple arithmetic relationships (subtrac tions,multiplications, additions), can be made in any known manner, eitherdigitally or in analogue form. FIGS. 12a and 12b are the basic diagrammost generally employed, of an ILM radar system in accordance with theinvention.

For the sake of clarity, the figure representing the whole systemaccording to the invention has been split into two parts, respectively12a and 12b the part 120 including conventional ILM radar receivercircuits, flight parameters data inputs, and angle computing circuits,and part 12 b including essentially the circuits improved according tothe invention, some of the conventional parts of the system appearing onpart 12b for the sake of clarity.

The system according to the invention comprises a certain number ofinputs generally speaking there are three groups of inputs to which areapplied input signals either supplied by other on board instruments, orissued from maps or charts available to the pilot and relating to theconsidered runway.

i. The group of flight parameters inputs the supply of these parameteris entirely conventional aboard modern commercial aircraft (aerodynamicand/or gyroscopic and/or inertia systems are used). The parameters usedhere are a. the roll angle 4) b. the pitchangle 6 c. the yaw angle AC d.the drift angle d The three first mentioned angular parameters arerelated to the aircraft attitude and are applied at input's I 1,, and Lythese three inputs are repeated in group at I for the sake of clarity.

The drift angle parameter input is shown at I (FIG. 12a).

ii. The group of the runway parameter inputs for introducing in thecomputing circuit values representing the characteristics of the runway.The values vary from one runway to the next and are generally manuallyentered into the circuits by the pilot or the co-pilot which is suppliedwith charts of the latters. They are the lengh L, the threshold length fand the width 2 m of the runway, shown at FIG. 18. These parameters areavailable at the general input I (FIG. 12,)

iii. The group of the reflector parameters, namely parameters a, b and eshown FIGS. 10 and 18. These parameters are available at input I,, FIGS.12a and 12b.

The form under which the input signals are available on board and inwhich they are supplied to the various inputs depends essentially on thekind of circuits to which they are applied (digital or analog circuits,and, if analog, dc or ac circuits).

This point will not be discussed here in detail, it being now entirelyconventional for those skilled in the art to convert optical signalsinto electrical, ac signals into dc signals, voltage signals toamplitude signals, pulse signals to continuous signal, digital signalsto analog and vice-versa.

The inputs having been defined, the system will now be described Itcomprises an antenna 1 which can operate as an elevation monopulsereceiver antenna, a scanner 2 for controlling the displacement of theantenna in azimuth at any one of two displacement speeds with any one oftwo amplitudes respectively associated therewith the displacement of theantenna being stabilised against rolling and pitching in a conventionalway, by a stabiliser device 3 whose control inputs are I and 1,, Theantenna is illustrated symbolically as having two radiator elements,corresponding to the said monopulse function, butthis does not mean thatit is necessarily geometrically formed of two elements as alreadystated, those skilled in the art will appreciate that a single radiatorsource of the multimode type can produce sum and difference signalssimultaneously.

The antenna 1 has a transmit-receive channel or sum channel, 1 1, and areception-only channel or difference channel 12. The channel 11 iscoupled to'a transmitreceive switch 4, for example an UHF circulator,and this, possibly through a polarisation switch 5, as is the case withmost weather radars.

The transmit terminal of the switch 4 is coupled to a transmitter 6 thelatter supplies UHF carrier pulses at a carrier frequency of 104 Mc/secfor example and at a repetition fr e quency F12 for example 200 c/s orFR (for example 2500 c/s, possibly wobbulated) of duration 1', (forexample 5 as) in the case of F and T (for example 0.4 #8) in the case ofF to this end the transmitter has a pulse width control input 61, thecontrol being manual and/or automatic, if the aircraft is equipped witha tracking computer which produces a signal for the input 61 when thecruising phase of the flight terminates, and also comprises an input 62controlling the repetition frequency, the two inputs being synchronisedmanually or automatically, as symbolically shown by the dashed linesbetween terminals 61 and 62.

Such antenna systems and transmitters are known and described forexample in the following documents German Pat. No. l, 234, 811; US. Pat.No. 3, 146, 448; US. Pat. No. 3, 177, 484; and French Patent No. 950,799.

The receiver terminal of the switch 4 is coupled to a mixer 7 eitherdirectly (weather radar operation), or, for the landing phase, through ahigh-pass UHF filter 8, whose cut-off frequency lies between f and f =f,A f; a switch 9 is provided for this purpose.

The mixer 7 is coupled to the output of a first oscillator 10 offrequency f, F,, F, being the intermediate frequency of the weatherradar receiver system. This oscillator is provided with automaticcorrection of the frequency f through a conventional feedback loopcomprising a mixer 111 coupled to outputs of the transmitter and of thelocal oscillator 10, an amplifier 112 and a frequency-discriminator 13centered on the frequency F,, for example 30 Mc/s The output of themixer 7 is coupled to the input of a switch 14 with two outputs 141 and142, the output 141 being coupled to a conventional weather radarreceiver 15, which, in the conventional way, comprises an ordinary videochannel, an isocontour channel (this, in particular, being equipped withmeans for controlling the gain as a function of the target range) and aPPI display. The setting of the switch 14 is carried out synchronouslywith the control, at 61, of the transmitter, the switch being inposition 141 when the transmitter is supplying pulses of longer duration7, and in position 142 when the transmitter is supplying pulses ofshorter duration, 1'

The terminal 142 is coupled to a second mixer 16 which is supplied onthe other hand with the output signal from a second local oscillator 17,of frequency 215 Mc/s for example. The output signal from this mixer, of

frequency 45 Mc/s in this example, afteramplification at 18 anddetection at 19, forms the video sum signal; preferably a known kind ofdual automatic gain control, is carried out in device 20, operating as afunction of the mean noise level and in sensitivity time control device21, as a function of the target range.

In the presence of other aircraft on final approach and equipped withthe same system of approach radar operating at substantially the samefrequency, the responses from targets to the pulses coming from one andthe same aircraft will be identified in the conventional way bycorrelation, for example correlation of three successive echoes,at 22the output signals from correlator 22 can then, likewise in knownfashion, be used to control the opening of range gates which, when thereceiver has locked onto the target echoes, will ensure that thesetargets are range tracked.

These signals, too, serve in the conventional way to determine theranges r r r between the aircraft and the reflectors to this end, theyare applied to a range computer 23 synchronised by the general pulseradar synchronising device,24, which of course likewise controls thetransmitter and the weather radar receiver 15.

The azimuth angles g g g are derived directly from a position detector25 which senses the antenna position and is coupled to the scanner 2.

These data as well as the range data are supplied to an arithmeticalcomputer 26 which is coupled to the yaw input 1,, and is supplied on theother hand with the elevational data AS AS AS supplied in the mannerdescribed hereinafter.

Let it be assumed that the computer is programmed for a given glideslope and a given runway configuration. In any other case, it comprisesI and I inputs by means of which are selected the glide slope p and theparameters L and b The elevation data is obtained from sum anddifference signals supplied by the monopulse antenna. To this end, thedifference channel 112 is coupled to a filter 81 identical to the filter8, coupled to a mixer 71 identical to the mixer 7 the output of themixer 71 is coupled to the mixer 161 identical to the mixer 16 Anoperator 50, the inputs of which are coupled to the outputs of themixers 16 and 161, forms the signals 2 +jA and 2 jA, the symbol jrepresenting a phase shift of 'rr/ 2 These signals are amplified andlimited, respectively, at 27 and 28, and the amplitude-phase detector 29coupled to the outputs of amplifier-limiters 27 and 28 produces theelevation deviations signal AS, (i l, 2, 3)which is supplied to theelevation input of thp computer 26 as shown.

This computer can be designed in a known manner: it will comprise acertain number of storage positions in which various values of thesignals ri (i l, 2 or 3), ASi and gi are recorded, the addressing of thecomputer store position being of course synchronised by the radar syncsystem (not shown for the sake of clarity), which supplies thetransmitter control pulses at 62. The computer further comprisesarithmetic operation circuits effecting the computation of AS,, a valueof the AS angle related to point B available at the output 265 (FIG.12b), the computation of r the distance of the aircraft to point B andthe computation of a coefficient K, which will be explained hereinafterand is necessary for the computation of angle y As it will be shown, allthese computations involve only pure arithmetic operations (addition,subtraction, multiplication and division) and are therefore computed bymeans of circuits readily obtained by those skilled in the art ofcomputers.

The distance r,, is substantially equal to where a and b are madeavailable to the computer at inputs I the coefi'rcient K being equal tothe second term, i.e.

The angle AS, may be taken within reasonable approximation equal to (A r(r AS r AS 1;.K where 1;, shown in FIG. 11, is a fixed value entered inthe computer circuit.

For the sake of clarity circuit 26 is shown again in FIG. 12b, where AS,r, g, represent globally the signals r AS g;

Each general input thus designated permits the display of severalparameters either in series or in parallel, each general inputrepresenting an assembly of terminals respectively corresponding to thevarious designated parameters.

The system further comprises five general computing units: The unit 201which determines the co-ordinates X Y Z of the aircraft relative to therunway trihedral the angle All] or decrabbing angle which is the angleby which it is necessary to rotate the aircraft about the normal axisfor making the aircraft longitudinal axis parallel to the runway axis atthe very instant when the wheels touch the ground the corrected values yand 8,. of the characteristic angles 'y and 8 the distance r e of theaircraft from the end of the runway opposed to the landing point Theunit 202 which determines the corrected values of the bearings of thereflectors and of the point B,,, which values are employed in the unit201;

The unit 203 or runway unit which determines, from the characteristicparameters of the runway and from those supplied by the foregoing units,the parameters necessary for a synthetic representation of the runway,which may possibly be associated with the representation of other flightparameters; and

the unit 204 for calculating the predicted stopping distance DS, and theanticollision unit 205.

The signal r available at the output 263 of computer 26 is subtractedfrom r;, in the subtractor 101 so as to furnish the value L representingthe distance B B Indeed the values a, b, and e are often data which areidentical from one runway to another, whereas the distance L can vary toa large extent (hundreds of meters or more) and it is of interest todetermine it from measurements rather than to display it.

In fact, its determination may constitute a test.

The signals e, of the input IB, and r of the input r, are subtracted inthe subtractor 102.

The output signal of the subtractor 102, which is equal to (r -e),represents the distance between the aircraft and the end of the runwayopposed to the landing point and is a precious indication not only ininitial and final approach but also during the running of the aircrafton the runway.

The co-ordinate X, of the aircraft may be considered equal to r, inabsolute value. The filtered value of X namely X is furnished at theoutput of the filter 104 which is connected in series with the inverter103 at the output 263.

The value of y is calculated from the corrected mean values G G G of thebearings g ,g ,g in response to n successive pulses upon each passage ofthe beam of the antenna respectively through the reflectors B, ,B ,B Afirst computer 105 to which are respectively applied the compensatedsignals g g determines the values and G V2 (G l-G b/4 (G -G The valuesof g g g supplied by detector 25 are compensated for the aircraftdisplacement in unit 310 which has inputs for g for parameters r a and b(input and for parameter C.

The computer 106 determines the signal The computer 107, having fourinputs respectively connected to the outputs of the computers 101, 105and 106 and to output 263 determines the corrected signal The circuits105, 106 and 310 constitute the unit 202. The co-ordinate Y, of theaircraft is determined in the multiplier 108 which receives the signalsr and respectively the filtered value, namely Y is furnished at theoutput of the low-pass filter 109 the lowpass filter 110, connected inseries with the output of the circuit 107, furnishes the filtered signalThe co-ordinate Z, of the aircraft is obtained by forming in themultiplier 111 the product B1,, which is thereafter filtered at 112. Thevalue of B is obtained in an arithmetic computer 26a having four inputsrespectively for signals AS (from circuit 29) for signal AS (fromcomputer 26) for signal r (from computer 23) and for signal L (fromcircuit 101) it elaborates signal B r lL (AS, A8 which is fed to themultiplier 111 and to the filter 113.

The signal B, filtered at 113, is applied to the subtractor 114 whichreceives also signal p and furnishes the filtered signal 8f equal to B pThe filtered signal at the output of filter 110, and G at the output offilter 115 and signal AC from the heading data generator arealgebraically added in the adder 116 which furnishes the signal A1]: 6;,-yc AC the decrabbing signal, since it represents the angular amplitudeof the yaw movement required when the wheels touch down, for making theaxis of the aircraft parallel to the runway axis and ensuring that theaircraft runs correctly along the ground. The circuits 101 to 104 and107 to 116 constitute the aforementioned unit 201 which furnishes thecharacteristic parameters of the position and of the attitude of theaircraft and precise values of the angular deviations 'y and 8 Thesignals r /L AS, A8 at the output of the circuit 26 a the signals G Gfrom the calculators and 106, the signal ,8; from the filter 113, thesignals 6 AC and 1 (input I from the gyroscopic unit of the aircraft(not shown) and the parameters defining the runway relative to thereflectors, are applied to the unit 203 which determines the parametersnecessary for the representation of the runway in perspective, whichrepresentation is known as a synthetic display in which the runwayappears as it would be viewed if visibility so permitted.

The principle of the representation of the runway is known.

It is described in patents filed by the Applicant and in particular inUS. Pat.'No. 3,486,010 and will not be described here, since the presentdevice has for its sole purpose to furnish the parameters necessary forthis representation.

The predicted stopping distance Ds is calculated in the unit 204connected to the output X of the unit 201 and to the accelerometer or tothe inertia unit of the aircraft furnishing the acceleration Xo along Bx,

The anticollision function is performed by the unit 205 coupled to theinputs r 17 and AS.

The various elements of the unit 201 are essentially conventionalelements algebric adders (101, 102, 116), inverter (103), multiplier(108, 1 l1), low-pass filters (104, 109, 110, 112) or conventionalcombinations of such elements (calculator 107).

The bearings 3 g g are of course measured at different instants further,the axis of the aircraft and consequently that of the radiation patternof the antenna is not strictly stable. The calculations of G and All:involve the bearings of the three reflectors and it is therefore wellnecessary to compensate for the effect of the displacement of theaircraft between the measurements and for the effect of possible yawmovements on the computed Go and All! values.

FIG. 13 shows in detail the bearing correcting circuit 310. The positiondetector 25 indicates a value 3,- (equal to g g g depending on theilluminated reflector), at the instant of measurement (the transmissionand reception instants are made to coincide since the speed of theaircraft, however high it may be, is at the most equal to a fewmillionths of the speed of propagation of electromagnetic waves).

in which j =1 ,2 n, an integrator 304 which calculates In fact the cours e of the aircraft at the moment of measuring is C C AC the value AC iscalculated from C in the computer 303 which forms the mean and thedifference C-C the circuit 303 being constructed very simply for exampleby means of a resistor and a capacitor if the signal C is a continuoussignal the signal AC is added to g,- in the algebraic adder 301. If themean course or heading of the aircraft relative to the runway All! isconstant or varies slowly andif the aircraft is near to the axis of therunway, the correction of G, as a function of the displacement of theaircraft can be achieved while assuming the derivatives of G, and of G,-to be equal G'i (FIG.14), being the bearing of the reflector Bi relativeto the axis of the runway, assuming that the aircraft is in the verticalplane of this axis (Y 0) this hearing G'i being calculated from r b anda in the computer circuit 305 by the relations G r 9* arc tan a/r,,b/2 Ga arc tan a/r,,+b/2 The differentiator 306 furnishes dG /dt which isadded at 307 to the mean value E For i= 3, G G and there is nocorrection These dynamic corrections are exact only so long as theaircraft A is in the vertical plane of the axis of the runway and solong as All; is substantially constant.

Now, All! is determined essentially by the component of the wind alongB,,Y,, and varies in the course of the descent with the altitude of theaircraft, which results in a residual error which is, however, small inthe final landing stage. The block-diagram of the unit 202 determiningthe bearings therefore includes as shown in FIG.

three correcting circuits 401, 402, 403, each of which has a loopidentical to the loop 300,

two differentiator and computer circuits, 405 and 406 for i l and 2,each of which being similar to circuits 305 and 306 of FIG. 13,

an arithmetic computer 408 coupled at the outputs of the G and Gcorrecting circuits and effecting the computation The synthetic displayof the runway can either be an image projected to infinity parallel tothe direction in in which the pilot is supposed to look (namely thedirection in which he would look at the runway if it was visible), thatis, an image which is superimposable on the runway head up display) oran image formed on an indicator placed on the instrument panel head downdisplay). In the latter case, the apparent diameter of the image of therunway is determined by the available space.

In both cases, the parameters necessary for the presentation of therunway are determined in the unit 203 from the angular and distance datadefining the altitude of the aircraft, its position relative to the axisof the runway defined by the points B B and from the geometric data ofthe runway (position of B and 13;, relative to the four corners of therunway for example). The angles and distances in question are furnishedby the previously described circuits (bearings G 6 angles All; and ,8,angles of elevation AS and AS by the (for example gyroscopic) navigationunit (attitudes 6 and -r variation in the heading or course AC the datarelating to the geometric configuration of the runway are displayed, forexample manually, for each landing.

In addition to the synthetic display of the runway, the indicatorconnected to the outputs of the unit 203 can also represent in theconventional manner the axis of the runway, the horizon, the theoreticalpath AT, the ground velocity vector and generally any other parameteravailable by other flight and navigation aids.

However, the number of parameters represented on this indicator will belimited so as to avoid an excessive complication of the display of therunway.

FIG. 18 shows the runway and the projection of the aircraft on theground V is the projection of the speed of the aircraft relative to theground on the horizontal plane.

The runway can be defined by the co-ordinates of the four points P P P Prelative to the aircraft trihedral.

The relative runway-reflectors arrangement is completely defined (FIGS.14 and 18) by the parameters e (distance between 8;, and the line P P L(distance B B 2a, b (projection of B B parallel to the axis of therunway) f (distance between B and the line P P and 2m the width of therunway.

FIG. 16 is a block diagram of an embodiment of the circuits 203 fordetermining the parameters of the runway display and FIG. 17 representsthe synthetic runway. The axes AY and AZ are the body axes (i.e. fixedin the aircraft); they are not necessarily displayed (in particular if ahead-up representation is used) but they have been included in theFigure so as to facilitate the explanation of the construction thereof.

The angles of elevation are low when the aircraft is approaching therunway (at the most a few degrees) and the apparent diameters arecalculated in the only horizontal plane).

It will be recalled that the visualisation is achieved by means ofcathode-ray tubes and that the different deviations along AY and AZ areobtained by applying voltages proportional to the desired deviations tothe deflection plates of the tube.

Hereinafter, in order to avoid overloading the description excessively,the same expressions will be employed indifferently for designating thevalues of the co-ordinates of a point of the synthetic runway and thevoltages, proportional to these co-ordinates, applied to the tubes.

The unit 203 comprises a first group of inputs connected to the devices(for example a gyroscopic unit) furnishing the angles AC 0 and I (inputI a second group of inputs connected to the outputs G ,AzIJ,-y,p ,B,(AS,,AS andX,,, Y Z ofthe devices 201, 202 and a third group of inputsfor displaying the characteristics of the runway parameters L, e,m,f forexample as indicated in FIG. 16 (input IRW) the characteristic L caneither be displayed manually or obtained automatically from the outputof the device 101 of the unit 201. In the absence of rolling, thehorizon would be represented by the dotted line H of the ordinate 0(FIG. 17).

The axis of the runway intersects, at point B having the ordinate Y All:(point at infinity on the axis) this line and the line H representingthe real horizon which makes the angle 1 with the line H The inputs 6and All; directly furnish the signals necessary for the representationof the line H and of the point B Hereinafter the co-ordinates along AYand AZ of each represented point will be marked by the indices Y AND Z.

The point B, is viewed (FIGS. 10 and 11) at angle 6 6 along AZ and atangle G, along AY An adder 501 (FIG.16) connected to the inputs and Bfurnishes B The reflector B is defined on axis B B, by Bay B The adder502 furnishes B An arithmetic computer 503 defines the co-ordinates ofthe points P P P P, relative to the runway trihedral B X Y Z, (FIG.

The co-ordinates along AY and AZ in the aircraft trihedral A X Y Z aredetermined, from the co-ordinates related to the runway trihedral, inthe co-ordinates transformer 504, which comprises in the known manner anassembly of multiplying and algebraic adding circuits, the co-ordinatesX,-Y,Z,- of a point P, (i 1, 2, 3, 4) being obtained from thecoordinates X Y Z by the relations of the following form H 0 and for thepoint B in which a a a b b b c c 0 are the linear functions of the sineand cosine of the angles 1 0 and All: defining the relative directionsof the two axes of the two trihedrals. The co-ordinates necessary forthe representation of the runway are the apparent diameters, that is theratig of the coordin a tes along AX and AY of the vectors AP, AP, AP AP,to the distances AP, AP AR, and AP, respectively.

A computer 505 determines these distances APi from the co-ordinates ofthe vectors in accordances with the relations (AB) AP AP AP where il,2,3,4 and AP AP AP are the co-ordinates of AI in the trihedral A X Y ZThe co-ordinates AP and AP are divided by AP, in the computer 506 whichfurnishes P and P Apart from the runway and the horizon, it is ofutility to represent also the previously-defined point B and the point Twhere the line parallell to B, T (FIG. 11) passing through A, intersectsthe runway. The position of this point relative to the runway enablesascertaining at each instant the location of the impact point if thedescent slope is maintained equal to p, assuming that the elevation andbearing corrections are insufficient.

AT, being parallel to B T, the co-ordinates of the point T along AY andAZ are respectively equal to All; and 0 p An adder 507 is connected tothe inputs p and 0 for this purpose. If the pilot does not modify thedirection of the aircraft velocity, the point of impact, starting fromthe actual position, would be at B,i.e. at the point of intersectionwith the ground of the line parallel to the velocity vector and passingthrough A. The co-ordinates of the point B are in which V and V are theco-ordinates of the ground speed and V its modulus, V representing thedrift angle of the aircraft. The values V Vy V are furnished to theinputs of the unit 203 by an auxiliary speed measuring system (Dopplerradar for example). In the absence of such a system, they can beobtained from the co-ordinates of the aircraft relative to the ground byderivation and filtering in this case, a circuit 508 is placed betweenthe speed inputs of the unit 203 and the outputs X, Y, Z, of the unit201. It comprises essentially three differentiators 701, 702, 703followed by three filters 704, 705, 706 the coordinates obtained at theoutput of the filters are connected to an assembly of circuits 804,805and 806 respectively identical to the circuits 504, 505 and 506 the samecircuits could possibly be employed with a switch.

The landing will be effected by bringing the points B, and B to T Thevarious output parameters of unit 203 are connected to the plates of thecathode ray devices employed for the direct or projected representationof the runway.

The parameters are represented in a discontinuous and recurrent mannerwith a sufficiently high repetition frequency to that the observer seesthem permanently. For this purpose, a switch 510, controlled cyclicallyby a clock 511, connects the outputs of the unit in succession to thedeflection plates of the tube CRT. A memory (not shown) may be placedbetween the switch 510 and the cathode ray tube, the read-out frequencybeing not necessarily the same as the writing frequency. Thesearrangements are known and will not be described in more detail here,since the object of the invention does not concern the production in theunit 203 of the representation of the runway which is known in the art,but the utilization, for this representation, of the precise andcomplete data produced by the units 201 and 202.

The predicted stopping distance is determined while the aircraft isrunning along the ground in the unit 204 by the relation in which X,,, VI" are respectively the coordinates of the aircraft, of its speed and ofits acceleration in the direction parallel to B.,X,,. An accelerometeror an inertia unit generally furnishes the coordinate X, The coordinateV can be obtained either by integration of IT, or by derivation from X0.or by means of an auxiliary Doppler radar. Generally, the co-ordinatesof the speed and acceleration are not available with respect to the axesB X B Y Thus the unit 204 shown in FIG. 19 comprises, in addition to anarithmetic computer 901 (multiplier, divider, and subtractor), acoordinate-transformer 902 which calculates from any coordinates of thespeed and acceleration the co-ordinates V and I;

Except during the landing phase of the flight, the data furnished by thesystem just described is also utilized in the unit 205 to precludecollision risk.

An obstacle M is dangerous if it is on or above the predicted path ofthe aircraft, as indicated in FIG. 20 in which AFa represents the axisof the beam of the elevation antenna, and AS the pointing error inelevation of a target M located at the distance D The unit 205calculates the expression E, D sin (1; AS I-I,, in which H is apredetermined safety altitude the output of the unit 205 is connected tothe P.P.I. indicator of the system so that only echoes in respect ofwhich E 0 are displayed. It is recalled that the ILM radar describedwith reference to FIGS. 12a and 12b have both meteo and landing aidfunctions and that, when operating for the meteo, the sum receptionchannel is connected to a conventional meteo receiver having a P.P.I.indicator.

In order to ensure the anti-collision function, as indicated in FIG. 21there is associated with the difference or A channel of the circuitdisclosed in FIG. 12a a switch 9,, and a switch 14,, whose types andfunctions are similar to those of the elements 9 and 14 of the sumchannel 2 this enables having the pointing error signal AS at the outputof the phase detector 29, also when the transmitter operates fordetecting the meterological disturbances.

The distance computer 23 automatically furnishes the distance D thecomputer 800 determines the signal E the calculations carried out by theunit 800 are conventional and can be effected in various known mannersfor example a subtractor (not shown) furnishes (1; AS) this differenceis applied to the first input of a multiplier (not shown) whose secondinput receives a voltage proportional to D the voltage H is subtractedfrom the output voltage of the multiplier. Aswitch 801 is placed betweenthe output 141. of the switch 14 and the meteo receiver 15. The switchis actuated by the output signal of the sign detector 802 connected tothe output of the unit 800 so as to connect the receiver 15 to thechannel only when E,, is negative.

For the sake of clarity of the figure, there have been shown in FIG. 21,among the elements of FIG. 12a, only those necessary for the descriptionof the anticollision device, the references employed for these elementsbeing the same as those used in FIG. 12a. In particular, there have notbeen shown the various switches actuated by the general synchronizingdevice of the radar system which ensures the correct connection of the ginput with the devices 401, 402, 403 in successlon Also are not shownthe details of the various eomputers which are employed in the circuitfor effecting successions of simple algebraic or trigonometricoperations and are in any case constructed in a manner known in the art.

The invention is of course not limited to the embodiment shown anddescribed.

We claim:

1. An Independent Landing Monitoring (ILM) pulse radar system for use incooperation with three reflectors B B2 and B (FIGS. l0, l4) and withauxiliary on board systems supplying flight parameters, said ILM systemsupplying bearing and elevation information B FIGS.10 and 11) relativeto the position of the aircraft on which it is mounted with respect tothe runway on which the aircraft has to land, similar to those suppliedby conventional I.L.S. Systems, said radar system further comprisingcompensation circuit means (310 ,FIG. 12b)for supplying bearing valuescompensated for the dynamic errors in the computation of said bearinginformation said circuit means comprising first correcting means (305,306) for compensating for the longitudinal displacement of the aircraftsaid first means having inputs for the reflector parameters (I and anoutput and second correcting means (300 and 303) for compensating forthe aircraft yaw movement said second means having an input for thecourse of the aircraft (C), an input for the antenna bearing indication(gi),an input coupled to said first correcting means output, and anoutput supplying said compensated value.

2. An ILM System according to claim 1, wherein said second correctingmeans comprise a course variation computer (303) having an input forsaid course (C) and an output (AC) a first algebraic adder (301, FIG.13)having a first addition input for said antenna bearing indication input,a second addition input ,coupled to said course variation computeroutput a sub traction input, and an output, averaging means (302) havingan input coupled to said adder output and an output, a second adder(307) having two inputs respectively coupled to said integrating meansoutput and to said first correcting means output, and an output, anintegrator means (304) having an input coupled to said second adderoutput, and an output coupled to said subtraction input, said integratoroutput supplying the compensated bearing value.

3. An ILM System according to claim 1, wherein said first correctingmeans comprise a computing means (305, FIG. 13) having an input for saidreflector parameters and an output, and differentiating means (306)having an input coupled to said last mentioned output.

4. An ILM System according to claim 1, comprising distance parameteroutputs (r r r r FIG. 12a) bearing indication outputs (G G elevationindication outputs (B) said ILM system further comprising furthercircuit means (201, FIG. 12 b) having inputs for said distanceparameters, for said bearing and elevation indication, and for saidreflector (1 parameters, said further circuit means (201) supplying thedistance of the aircraft to the far end of the runway (r the coordinatesof said aircraft X,,,, Y Z,,,) with respect to a trihedral fixed to therunway of the decrabing angle (Art), and the ILS angular deviation (7and 8,

5. An ILM system according to claim 4, having an input for the aircraftacceleration parallel to the 'runway axis F FIG. 12b said system furtherIcomprising distance computing means (204) having an input for saidaircraft abcissa in said trihedral (X an input for said acceleration andan output supplying the predicted stopping distance of the aircraft.

6. An ILM system according to claim 1 having distance (r) antenna axiselevation (1;) and elevation error signal (AS) outputs, said systemfurther comprising an anti-collision circuit (205 FIG. 12b) havingrespective inputs coupled to said last mentioned outputs.

7. An ILM system according to claim 1 having an input for the aircraftattitudes (IAT) an input for the runway parameters (IRW) associated witha runway display means having display inputs said ILM system furthercomprising further circuit means (203) FIG. 12 b) for supplying saidrunway display means, said last mentioned further circuit means havingrespective inputs for said aircraft attitudes, for said runwayparameters and for said bearing values supplied by said compensationcircuit means, and respective outputs coupled respectively to saidrunway display means inputs.

1. An Independent Landing Monitoring (ILM) pulse radar system for use incooperation with three reflectors B1 , B2 and B3 (FIGS. 10, 14) and withauxiliary on board - systems supplying flight parameters, said ILMsystem supplying bearing and elevation information ( gamma , deltaFIGS.10 and 11) relative to the position of the aircraft on which it ismounted with respect to the runway on which the aircraft has to land,similar to those supplied by conventional I.L.S. Systems, said radarsystem further comprising : compensation circuit means (310 ,FIG.12b)for supplying bearing values compensated for the dynamic errors inthe computation of said bearing information said circuit meanscomprising : first correcting means (305, 306) for compensating for thelongitudinal displacement of the aircraft , said first means havinginputs for the reflector parameters (IB ) and an output , and secondcorrecting means (300 and 303) for compensating for the aircraft yawmovement , said second means having an input for the course of theaircraft (C), an input for the antenna bearing indication (gi),an inputcoupled to said first correcting means output, and an output supplyingsaid compensated value.
 2. An ILM System according to claim 1, whereinsaid second correcting means comprise a course variation computer (303)having an input for said course (C) and an output ( Delta C) , a firstalgebraic adder (301, FIG.13) having a first addition input for saidantenna bearing indication input, a second addition input ,coupled tosaid course variation computer output , a subtraction input, and anoutput, averaging means (302) having an input coupled to said adderoutput , and an output, a second adder (307) having two inputsrespectively coupled to said integrating means output and to said firstcorrecting means output, and an output, an integrator means (304) havingan input coupled to said second adder output, and an output coupled tosaid subtraction input, said integrator output supplying the compensatedbearing value.
 3. An ILM System according to claim 1, wherein said firstcorrecting means comprise a computing means (305, FIG. 13) having aninput . . . for said reflector parameters and an output, anddifferentiating means (306) having an input coupled to said lastmentioned output.
 4. An ILM System according to claim 1, comprisingdistance parameter outputs (ro, r1 , r2 , r3 FIG. 12a) , bearingindication outputs (Go, G3), elevation indication outputs ( Beta ) ,said ILM system further comprising further circuit means (201, FIG. 12b) having inputs for said distance parameters, for said bearing andelevation indication, and for said reflector (IB) parameters, saidfurther circuit means (201) . . . supplying the distance of the aircraftto the far end of the runway (r3 e) , the coordinates of said aircraft (Xof, Yof, Zof) with respect to a trihedral fixed to the runway of thedecrabing angle ( Delta psi ), and the ILS angular deviation ( gamma cfand delta f ) .
 5. An ILM system according to claim 4, having an inputfor the aircraft acceleration parallel to the runway axis ( < xo FIG.12b ) , said system further comprising distance computing means (204)having an input for said aircraft abcissa in said trihedral (Xof) , aninput for said acceleration and an output supplying the predictedstopping distance of the aircraft.
 6. An ILM system according to claim 1having distance (r) antenna axis elevation ( eta ) and elevation errorsignal ( Delta S) outputs, said system further comprising ananti-collision circuit (205 FIG. 12b) having respective inputs coupledto said last mentioned outputs.
 7. An ILM . . . system according toclaim 1 having an input for the aircraft attitudes (IAT) , an input forthe runway parameters (IRW) associated with a runway display meanshaving display inputs said ILM system further comprising further circuitmeans (203) FIG. 12 b) for supplying said runway display means, saidlast mentioned further circuit means having respective inputs for saidaircraft attitudes, for said runway parameters and for said bearingvalues supplied by said compensation circuit means, and respectiveoutputs coupled respectively to said runway display means inputs.
 8. AnILM System according to claim 1 transmitting pulses of a first givenduration at a first given repetition frequency, comprising a monopulseantenna, a difference reception channel ( Delta ) and a sun receptionchannel, said sun channel comprising frequency changing means (7FIG.12a) having an output, said ILM system further comprising means forweather radar operation, said last mentioned means comprising a weatherreceiver (15 FIG. 12 b and FIG.21), means for synchroneously (i)transmitting pulses of a second given duration longer than said firstone and at a second repetition frequency smaller than said first one,and means for (ii) coupling said changing means output to said weatherreceiver.